In many operator controlled systems, operator induced oscillations are sometimes experienced, particularly when very rapid corrections are made so that the probability and extent of overcorrection increases. For example, in an aircraft, particuarly one of high performance or of special flight characteristics, such as the space shuttle vehicle under development by NASA, pilot-induced oscillations (PIO) may be experienced during the flare and landing. The frequency of the PIO in the pitch axis for the space shuttle vehicle was approximately 3 rad/sec, and the pilot utilized a significant portion of his full control authority. Analysis has indicated that the source of the problem is a combination of basic vehicle handling qualities, time delay of approximately 0.04 sec through the digital flight control system computer, and rate limit of 20 deg/sec of the elevator actuators.
Each control system will have a characteristic control gearing schedule for an operator input that is generally not linear but rather curved in both quadrants, with a positive response for an operator input of one control direction, and a negative response for an operator input of the opposite direction, with the response increasing nonlinearly as the operator input increases in either direction. The gearing schedule may include a "deadband" near center to decrease the system sensitivity to small operator inputs, but beyond that "deadband," the response of the system to a control input can be quite large, resulting in the possibility of overcorrection leading to a large reverse correction, and then oscillation if the process is repeated. To prevent oscillation, it is desirable to reduce the operator's authority over the control system as the frequency of the operator's control input increases. Typical of the prior art is disclosed in a U.S. Pat. No. 4,030,011 titled Multimode Control Systems.
In that prior art system, a control input signal is simultaneously high-pass filtered and low-pass filtered. The high-pass filtered signal is amplitude limited and added to the low-pass filtered signal. The sum is then subtracted from the control input signal to form the basic system control signal to which an accelerometer signal similarly conditioned is added. A further input from a pitch rate gyro is conditioned and combined with the system control signal applied to an elevator servo of an aircraft. However, these accelerometer and pitch rate signals are independent of the problem of reducing the operator's authority as a function of the operator's control input frequency.
An advantage of the prior art system is that mode switches are not required, but the frequency response characteristic is basically one of very low gain for low frequency signals, with increasing gain to a limit as frequency increases. The limit to which gain increases with frequency is set by the high-pass filter limiter. The result is the ability to limit the pilot control signal of large amplitude and, to a lesser extent, limit the pilot control signal of high frequency content. That does not satisfy the requirement for the combination of high pilot gain during the critical landing task, coupled with the ability to reduce the rate of the control system for large pilot inputs. This requirement is a significant factor in pilot-induced oscillations. Linear filtering techniques of the type used in the aforesaid patent have not been able to resolve the PIO problem since the required rate reduction involves unacceptable phase lags.